This invention relates generally to gas turbine engine blades, and more specifically to methods and apparatus for facilitating preventing failure of gas turbine engine blades.
At least some known gas turbine engines include a core engine having, in serial flow arrangement, a fan assembly and a high pressure compressor which compress airflow entering the engine. A combustor ignites a fuel-air mixture which is then channeled through a turbine nozzle assembly towards low and high pressure turbines which each include a plurality of rotor blades that extract rotational energy from airflow exiting the combustor.
Failure of a component within a system may significantly damage the system and/or other components within the system, and may also require system operation be suspended while the failed component is replaced or repaired. More particularly, when the component is a turbofan gas turbine engine fan blade, a blade-out may cause damage to a blade that is downstream from the released blade. More specifically, depending upon the severity of the damage to the downstream blade, other blades downstream from the released blade or the damaged trailing blade may also be damaged. Damage to the trailing blade may cause the trailing blade to fail, thereby possibly requiring operation of the turbofan gas turbine engine be suspended, and/or damage to other fan blades and/or other components within the turbofan gas turbine engine.
For example, at least some known turbofan gas turbine engines include a fan base having a plurality of fan blades extending radially outwardly therefrom. The impact of a released blade upon a trailing blade may cause the trailing blade to rock about an axis tangential to rotation of the fan. The trailing blade initially rocks about the tangential axis toward a forward-section of the trailing blade such that the trailing blade may be dislodged radially outwardly away from its disk slot. The motion of the trailing blade about the tangential axis then reverses due to rotation of the fan, causing the trailing blade to rock backwards toward an aft end of the trailing blade. The rocking of the blade may induce compressive and tensile stresses in the blade. The magnitude of these tensile and compressive stresses in the trailing blade may exceed the failure threshold of the blade material causing the trailing blade to fail.
In one aspect, a method is provided for fabricating a fan assembly for a gas turbine engine. The method includes forming a blade including an airfoil extending from an integral dovetail used to mount the blade within the rotor assembly, and extending a projection from at least a portion of the blade, such that the stresses induced within at least a portion of the blade are facilitated to be maintained below a predetermined failure threshold for the blade to facilitate preventing failure of the blade.
In another aspect, a gas turbine engine blade is provided that includes an airfoil, a dovetail formed integrally with said airfoil, and a projection that extends outwardly from at least one of the airfoil and the dovetail. The projection is configured to facilitate at least partially restricting movement of the blade to facilitate preventing failure of the blade.
In yet another aspect, a fan assembly for a gas turbine engine is provided. The fan assembly includes a fan hub, and at least one fan blade that extends radially outwardly from the fan hub. The fan blade includes a dovetail, an airfoil extending outwardly from the dovetail, and a projection that extends outwardly from the dovetail for maintaining stress induced within at least one of the dovetail and the airfoil below a predetermined failure threshold for the fan blade.